1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically for an air cooled turbine blade with showerhead film cooling holes for cooling a leading edge surface.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
To cool the leading edge region of a rotor blade or a stator vane, an arrangement of film cooling holes and gill holes are used. Rotor blades differ from stator vanes in that the rotor blade is rotating which effects the ejection of any cooling air from the holes. FIG. 1 shows a cross section view of a leading edge (L/E) region of a rotor blade with three film cooling holes 11 located in the leading edge and two gill holes located on the pressure side (P/S) and the suction side (S/S) of the L/E film cooling holes. The middle film cooling hole is located at a stagnation line which is where the hot gas stream strikes the airfoil at 90 degrees to the surface. The other two film cooling holes are located adjacent to and on the P/S and the S/S of the stagnation film cooling hole. This arrangement is formed along the entire airfoil surface from the platform to the blade tip. The film holes 11 and the gill holes 12 are supplied with cooling air from a cooling air supply channel 13 through a row of metering and impingement holes 14 that open into a leading edge impingement cavity 15. The L/E impingement cavity 15 can be formed from one long cavity or several cavities that form individual and separated compartments for the purpose of customizing to cooling air flow and pressure into the respective cavity depending upon the heat load and external gas pressure along the L/E of the airfoil. FIG. 2 shows a cross section view of the entire blade with the L/E cooling circuit described in FIG. 1.
FIG. 3 shows a cross section side view of the film cooling hole 11 along the stagnation line through the line A-A in FIG. 1. The three rows of film cooling holes on the L/E are inclined at around 20 to 30 degrees from the L/E airfoil surface and towards the blade tip. FIG. 4 shows a front view of the L/E film cooling holes with the stagnation row in the middle and the P/S row on the right and the S/S row on the left. The main problem with this L/E film cooling hole design is the over-lapping of film cooling ejection flow in a rotational environment of the rotor blade and a lacking of film coverage for the blade L/E region. As seen in FIG. 4, film cooling air ejected from the middle row does not cover the entire surface around the P/S or S/S rows. There are areas that are not covered with a layer of film cooling air. As a result of this lack of full coverage, a hot streak is developed between film holes.
One disadvantage of the showerhead arrangement of the prior art of FIG. 4 is the use of film cooling holes with a constant spanwise angle without tailoring to the leading edge region heat load as well as mechanical loads. As a result of the prior art film cooling, an over-cooling occurs at the blade lower span and/or an over-stress occurs at the blade root section if a low surface angle cooling hole is used, such as a 20 degree angle film hole to the airfoil surface. If a high surface angle hole is used, such as a 35 degree film hole, then an under-cooling of the leading edge will occur in the high heat load region of the airfoil leading edge. FIG. 3 shows a graph of the blade heat load and FIG. 4 shows a graph of the mechanical loads on the blade leading edge region.